Satellite launch system

ABSTRACT

A system for launching aerospace payloads includes a wingless, unmanned modified lifting body spacecraft ( 100 ), with a payload compartment in the forward section of the spacecraft. The spacecraft is propelled by hybrid rockets clustered in the aft section of the spacecraft. Reaction control system (RCS) modules control the flight path and its associated avionics hardware and software. This system also includes a carrier aircraft ( 200 ) configured to air-launch the spacecraft. The carrier aircraft includes a flight operations control system, which monitors the spacecraft&#39;s payload and monitors and controls launch and flight operations of the spacecraft. A ground-based mission control system monitors and controls the spacecraft&#39;s payload and monitors and controls the launch and flight operations of the spacecraft.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.16/046,153, filed Jul. 26, 2018; which claims priority to U.S. patentapplication Ser. No. 14/934,148, filed Nov. 6, 2015; which claimspriority to Provisional Application No. 62/082,450, filed Nov. 20, 2014.the contents of each of which are incorporated herein by reference intheir entirety..

TECHNICAL FIELD

The present disclosure relates generally to satellite launch systems,and more particularly to a system for launching satellites (or “sats”)of various sizes, including small-sats, mini-sats, nano-sats, as well asother payloads that may be placed into space.

BACKGROUND

Satellites are essential for many aspects of modem life. GPS,television, broadcast, mobile communications devices, all rely on theability to place satellites in orbit.

Following advances in engineering and technology, particularly inminiaturization, the lack of an affordable, reliable and easilyaccessible launch service for small satellites has all but groundedflight-ready experiments and generally stifled progress in the field forseveral years. Having commissioned a study in summer, 2014, “Air Launchor Ground Launch: A Small Satellite Comparative Study”, to discover thereasons for this lapse, the inventors have undertaken to outline aspecific air launch response. The concept of air launching spacevehicles is well known through the NASA/Dryden B-52 at Edwards AFB inCalifornia, and the American space initiative largely owes its evolutionto the X-I5 and the Space Shuttle, both developed with data from airlaunched operations.

While avoiding many vagaries (uncertain weather, scheduling conflicts,flight irregularities, restrictive and expensive protocols, etc.)associated with ground based operations, the delays, high price tags andinsurance costs nevertheless remain problematic. We learned from thestudy that air launching has challenges of its own; the negative impacton the performance of the “pitch-up” maneuver immediately afterhorizontal separation is far from trivial. Essentially, on release, thevehicle develops negative vertical delta-V. However, the burn rate ofrocket fuels is very rapid, generally around one minute or a few secondslonger, therefore considerable first stage energy is depleted inregaining lost altitude and establishing a positive climb profile.Additionally, while rocket aerodynamics are very low drag, they are alsovery low lift.

SUMMARY

It is the intention of the inventors to utilize the information from thepreviously commissioned study to design and develop a system including aspacecraft and a carrier aircraft to air launch said spacecraftspecifically to enable the affordable and reliable launch of smallsatellites and other light aerospace payloads as a service to thesmall-sat industry.

The present invention provides a system for launching satellites,including small-sats, mini-sats, nano-sats, medical and scientificexperiments, suborbital, orbital and other aerospace payloads, whichincludes a modified and optimized existing carrier aircraft, astreamlined, unmanned, rocket-propelled lifting body spacecraft, airlaunched from said carrier aircraft and containing in addition to itsown propulsion, the payload, staging, and insertion rocketry necessaryto the mission and provisions for protecting such payload while loading,fueling, transit to and mating with the carrier aircraft, towing,taxiing, conventional takeoff from the ground, climb and cruise to theselected launch point (LP) and high altitude orbital injection, as wellas tracking, navigation and control hardware, software and otherequipment necessary to establish a safe, reliable and affordablesmall-sat delivery service.

The inventive system for launching aerospace payloads comprises awingless, unmanned modified lifting body spacecraft including a payloadcompartment in a forward section of the spacecraft, hybrid rocketsclustered in an aft section of the spacecraft, reaction control system(RCS) modules configured to control a flight path of the spacecraft, andassociated avionics hardware and software; and a carrier aircraftconfigured to air-launch the spacecraft.

An illustrative embodiment further comprises a flight operations controlsystem, carried in the carrier aircraft, configured to monitor thespacecraft's payload and to monitor and control launch and flightoperations of the spacecraft. The illustrative embodiment may alsoinclude a ground-based mission control system.

In the illustrative embodiment, the hybrid rockets are contained withina shell composed of composite panels forming the aerodynamic shell ofthe lifting body. Moreover, the composite panels separately encase aportion of the spacecraft housing the hybrid boosters and cover thepayload compartment. In the illustrative embodiment, the compositepanels join at a horizontal chine line, and are configured to bejettisoned with pyrotechnical charges that separate one or more panelsfrom the remaining structure of the spacecraft.

The illustrative embodiment further comprises control hardware andassociated software configured to activate the pyrotechnical charges forthe appropriate panels at the appropriate phase of the flight profile,such that panels that encase the hybrid boosters are released togetherand panels that encase the payload bay are released together.

The illustrative embodiment further comprises chines that widen aftforming wye-shaped stabilizers, wherein each arm of the wye is cantedoutboard from the vertical. A horizontal arm of the wye is fitted withelevons and/or speed breaks.

In the illustrative embodiment, the hybrid rockets comprise two Stage 01boosters, two Stage 1 boosters, and one Stage 2 booster. The two Stage01 boosters are smaller than the Stage 1 and Stage 2 boosters, and theStage 01 boosters are positioned outboard of their adjacent Stage 1booster, vertically centered on a horizontal plane of the spacecraft.Moreover, the Stage 01 boosters are mated to the remainder of thespacecraft so as to transfer their thrust to the entire spacecraft, andare configured to be jettisoned with pyrotechnical charges. Further, theStage 01 boosters are ignited first after the spacecraft separates fromthe carrier aircraft, and are configured to orient the spacecraft for aninitial boost phase of the flight profile. In the illustrativeembodiment, the Stage 1 boosters are positioned on either side of theStage 2 booster, inboard of their respective Stage 01 booster,vertically centered on a horizontal plane of the spacecraft. In thisembodiment, the Stage 1 boosters are mated to the spacecraft so as totransfer their thrust to the entire spacecraft, both before and afterjettisoning the Stage 01 boosters, and are configured to be jettisonedwith pyrotechnical charges. The Stage 1 boosters are ignited after thespacecraft separates from the carrier aircraft and after the ignition ofthe Stage 01 boosters, when the spacecraft is in the correct orientationto begin the initial boost phase of the flight profile. Moreover, theStage 2 booster is centrally positioned on horizontal and verticalplanes of the spacecraft, and the Stage 2 booster is mated to thespacecraft so as to transfer its thrust to the entire spacecraft, bothbefore and after jettisoning the Stage 1 boosters, and is configured tobe jettisoned with pyrotechnical charges.

In the illustrative embodiment, each hybrid rocket comprises apressurized oxidizer tank, the reaction chamber with solid fuel, andigniter, combustion channels, exhaust nozzle, valves to control oxidizerflow, and hardware and associated software to monitor and control theoperation of the booster. In this embodiment, each hybrid rocket furthercomprises an oxidizer tank pressurized so as to control the flow of theoxidizer into the reaction chamber in order to modulate thrust bystarting and controlling the rate of combustion; hardware and associatedsoftware that monitors and controls sensors and actuators that managethe oxidizer flow as well as proper temperature and pressure of theoxidizer within the tank. In addition, each hybrid rocket furthercomprises: a reaction chamber configured to contain fuel for the rocket,the igniter, and combustion channel where the fuel and oxidizer combineand are ignited to generate propulsive force; and ablative materialsused to maintain correct operational temperatures within the combustionchamber. Moreover, each hybrid rocket further comprises: an exhaustsystem assembly including an interface to the reaction chamber, throat,nozzle, sensors, actuators, hardware and associated software to monitorand control exhaust flow.

In the illustrative embodiment, each hybrid rocket further comprisespyrotechnical charges that detach the rocket from structural members ofthe spacecraft to which the hybrid rocket is mated, and hardware andassociated software that interface with the flight system avionics, totrigger the charges at the appropriate stage of the flight profile. Inthis embodiment, each hybrid rocket further comprises: a casing withassociated structural members and interfaces that aggregate thesubcomponents of the rocket into a single component, wherein the casing,via its associated structural members and interfaces, is connected withother components as required by the final spacecraft assembly.

In the illustrative embodiment, the RCS comprises four individualmodules, located symmetrically about horizontal and vertical planes ofthe spacecraft, forward of the center of gravity, forward of therespective Stage 01 boosters, mounting to the aft face of the payloadcompartment bulkhead. In this embodiment, each RCS module comprises apressured tank of monopropellent (e.g. concentrated hydrogen peroxide),reaction chambers containing a catalyst (e.g. Tungsten mesh), andmultiple exhaust ports, oriented to provide pitch, roll, yaw, andtranslation control of the spacecraft. Moreover, each RCS module furthercomprises hardware and associated software configured to monitor andcontrol sensors and actuators that manage the oxidizer flow and propertemperature and pressure of the oxidizer within the tank.

The illustrative embodiment further comprises a cowling and shieldingaround the exhaust ports to interface with and protect the panelscomprising an aerodynamic shell to allow the thrusters to operate whilethe sheathing is in place. This embodiment also comprises a propulsionmodule made up of an assemblage of the rockets, structural members, andavionics.

In the illustrative embodiment, the payload compartment comprises apayload compartment bulkhead; structural members to support compositepanels enclosing the payload bay until they are jettisoned; apparatusconfigured to secure and deploy payload components; a system of sensors,actuators, hardware and associated software configured to monitor thestate of the payload bay, the payload, and to control jettisoning of thecomposite panels at an appropriate stage of the flight profile.

In the illustrative embodiment, the payload compartment bulkhead isconfigured as a means by which the propulsion module transfers motiveforce to the payload; as a means by which the apparatus that secures anddeploys the payload components is secured to the spacecraft andinterfaced to the spacecraft avionics; and as a platform from which thepayload components, upon achieving orbit, are placed into the initialphase of their orbital insertion profile.

In the illustrative embodiment, the monitoring and control of thespacecraft, including the payload, is managed by hardware and associatedsoftware comprising a top-level avionics infrastructure for the entirespacecraft, the infrastructure interfacing with control systems of allspacecraft sub-components.

In the illustrative embodiment, the infrastructure accepts operationaldirectives for the spacecraft and coordinates the activity of systems torealize and apply the directive, and compiles and presents alloperational status for all spacecraft systems.

In the illustrative embodiment, the carrier aircraft comprises theairframe itself, configured to conduct an air launch of the spacecraft;a physical interface that mates the spacecraft to the carrier aircraft;engines that provide propulsion for the carrier aircraft; and apparatusto configure a portion of the space inside the fuselage to house aflight operations control system.

The illustrative embodiment further comprises a physical interface thatmates the spacecraft to the carrier aircraft, supports the spacecraftduring takeoff and flight to the launch point, and deploys thespacecraft at launch. In this embodiment, the physical interfaceattaches to the carrier aircraft by a set of hard points bound toload-bearing structural members of the carrier aircraft, the hard pointsbeing positioned to enforce stability during takeoff and flight. Thephysical interface also includes a decoupling apparatus where thespacecraft connects with the interface such that, at the time of launch,the interface releases the connection with all hard points on thespacecraft simultaneously.

In the illustrative embodiment, the flight operations control systemcomprises avionics hardware and associated software configured tomonitor the payload, and to monitor and control the spacecraft from thetime it is mated to the carrier aircraft until control is transferred tothe ground-based mission control system; communications apparatus; andpower supply and conditioning apparatus.

In the illustrative embodiment, the ground-based mission control systemcomprises hardware and associated software configured to monitor andcontrol the payload, to monitor and control the spacecraft over itsoperational lifetime; communications apparatus; and power supply andconditioning apparatus.

An additional important feature includes a low, wheeled, concave dollydesigned to facilitate servicing and loading of the spacecraft to thecarrier aircraft. Further, an illustrative best embodiment of theinvention includes a provision for truncating standard rocket nose conesat the payload bulkhead in favor of utilizing the said space craft'smuch larger magnum payload bay (MPB). A further illustrative embodimentinvolves configuring the spacecraft as a flying test bed for newspacecraft prototypes. Moreover, illustrative embodiments may include anoperations specifications and limitations storage medium (e.g., acomputer readable medium) describing in detail the manner of operatingthe launch system, together with checklists for each phase, includingstandard, irregular, and emergency procedures. Other aspects of thepresent invention are described below and depicted in the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an exterior plan view of the LB-1 spacecraft (100) of thepresent invention.

FIG. 2 is an exterior elevation view of the LB-1 spacecraft of thepresent invention.

FIG. 3 is a front exterior view of the LB-1 spacecraft and the carrieraircraft (200).

FIG. 4 is a detail plan view of the LB-1 spacecraft of the presentinvention.

FIG. 5 is a detail elevation view of the LB-1 spacecraft of the presentinvention.

FIG. 6 is a section at line A-A of the LB-1 spacecraft of the presentinvention.

FIG. 7 is a plan view of conventional booster detail of the LB-1 of thepresent invention.

FIG. 8 is a plan view of positioning and attachment of said LB-1spacecraft of the present invention.

DETAILED DESCRIPTION OF ILLUSTRATIVE EMBODIMENTS

All illustrations of the drawings are for the purpose of describingillustrative embodiments and are not intended to limit the scope of thepresent invention.

The present invention provides a system for launching satellites,including small-sats, mini-sats, nano-sats, medical and scientificexperiments, suborbital, orbital and other aerospace payloads, includinga modified and optimized existing carrier aircraft, a streamlined,unmanned rocket-propelled lifting body spacecraft (100), air launchedfrom said carrier aircraft (200) and containing in addition to its ownpropulsion, the payload, staging, propulsion and insertion rocketrynecessary to the mission and the provisions for protecting said payloadduring loading, fueling, transit to and mating with said carrieraircraft, towing, taxiing, conventional takeoff from the runway, climband cruise to the selected launch point (LP) and high altitude release,as well as the tracking, navigation and control hardware, software andother equipment to effect a safe, reliable and affordable deliveryservice, including:

FIGS. 1, 2, 3 depict exterior views of LB-1, an unmanned,rocket-powered, wingless lifting body spacecraft 100, assembled ofcommercially available composite rocket boosters, complete with solid orhybrid fueled motors and strapping hardware. The spacecraft's liftingbody characteristics are designed to mitigate lift and altitude lossesat the horizontal release maneuver (HRM) and its wingless profile allowsattachment between the engines and landing gear of the carrier aircraft200 (as shown in FIG. 3, at the lower fuselage 7). Joined at chine lines1A & B, streamlined airfoils of carbon/composite form the nose cone, 2A& 2B, and main body fairings 3A & 3B to create a strong, lightweightcovering and provide a lift factor of approximately 65 pounds per squarefoot. Reference number 3AA (FIG. 3) illustrates an attachment fairing.Mid-body horizontal chines 1A & 1B on each side gradually widen aft ofthe nose cone from 2 to 3 feet, terminating in upper and lower “wye”stabilizers 4A & 4B canted outboard 60 degrees from the horizontal andfitted with split elevons 5A & 5B to maintain roll control in theatmosphere. Additionally, the aft chines are fitted with splithorizontal elevons for pitch control and use as speed brakes.

The spacecraft's body cross section may be described as a flattenedellipse with a longitudinally placed, laterally centered conventional2/3-stage rocket booster flanked by symmetrical pairs of propellantboosters of decreasing diameters and a wide, tapering nose cone toestablish the desired cross-sectional airfoil.

Four thrusters 6A & 6B (FIGS. 3 and 5) have been provided near theforward end of the upper and lower outboard boosters to increasestability during pitch-up. To avoid waste of Stage 1 thrust, smalloutboard boosters designated “Stage 01” will be ignited to accomplishthe pitch-up maneuver prior to Stage 1 ignition. (See FIGS. 4 and 7.) Itis foreseen that this combination along with the aforesaid improvementin lift will result in a considerably smoother, more controlled andeconomical spacecraft rotation.

FIGS. 4, 5, 6: Carrier aircraft belly vertical clearance 7 and landinggear fore and aft clearance 7A & 7B are shown with LB-1 mounted.Reference number 7C indicates the coupling keel and 7D the mountingsnubbers. The payload bay is at 9, and potential carrier aircraft hardpoint connections at 10. (See FIG. 4.) The arrangement of boosterspermitting maximum opportunity to accommodate various loading optionsand combinations of payload types is shown in relation to stages. Thepreferred embodiment provides that the rocket casings at 11 (FIG. 6) maybe truncated at the firewall along section line A-A and the entire noseof the spacecraft or selected portions thereof may be utilized.

FIG. 7 depicts a plan view showing a conventional booster display at 13.To enable and control the cost of this “quick-change” facility it isplanned that several firewall/payload plate options will be madeavailable at loading sites.

FIG. 8 provides a positioning diagram for mating of the LB-1 utilizingthe carrier aircraft hard point connections 10. An outline of thetransport dolly chassis 15 demonstrates lead-in guidance and criticalcomponent clearances. A low, wheeled concave dolly shaped to accept,center and support the convex lower spacecraft half for preciselyplacing the propellant boosters and to support the same during transit,servicing, fueling, applying the upper spacecraft half and towing underthe carrier aircraft for mounting and supplying battery power to saidspacecraft components.

Facilities for the monitoring and audible alarm of latching/sealingmechanisms, rising temperatures, leakage of oxidizer, suppression offire and other safety measures which may be provided at the spacecraft,and the carrier aircraft cockpit and launch control stations, separatefrom similar systems in the carrier aircraft.

Attachments and adaptors on the carrier aircraft and the spacecraft toenable the quick attachment/release of the spacecraft may also beprovided.

Facilities in the carrier aircraft and on the ground to remotely controlthe spacecraft as a mission-abort/reentry vehicle may also be provided.

Computerized Operations Specifications and irregular and emergencyprocedures and checklists to be performed by crew members will govern inall phases of the mission.

The LB-1 spacecraft is scalable over the range of potential carrieraircraft to suit the requirements of smaller or larger payloads.

The LB-1 is designed for polar and equatorial launch missions.

Although the invention has been described in terms of its preferredembodiment, it is to be understood that many other possiblemodifications and variations can be made without departing from thespirit and scope of the invention. Launch preparations, includingassembling, loading, and attaching the LB-I to the aircraft, include, inthe following order:

-   -   1. Lower body fairing, chines, stabs.    -   2. Thrust plate.    -   3. Firewall    -   4. Stage 2/3 booster.    -   5. Stage I boosters & straps.    -   6. Stage OI boosters & straps.    -   7. Left & Right Thrusters.    -   8. Payload plate.    -   9. Upper Body fairing, chines, stabs.    -   10. Secure cargo in Payload Bay.    -   11. Place nose cone, secure & check all fasteners.    -   12. Align LB-I beneath carrier aircraft.    -   13. Complete LB-I attachment checklist.    -   14. Attach LB-I to carrier aircraft & secure.

Flight operations, including towing, taxiing, takeoff, and flight,require the following:

-   -   1. All towing operations with LB-I attached shall be conducted        in radio contact with qualified ground crew ahead and behind the        carrier aircraft and ground level visibility of at least 3        nautical miles.    -   2. Prior to engine start all landing gear and tires shall be        checked for damage or irregularities and the captain advised.    -   3. Immediately prior to every take-off with LB-1 attached the        ground crew shall scan the takeoff runway for foreign objects        and remove any debris advising the captain by radio that the        runway surface is safe for takeoff.    -   4. When the captain receives the ground crew “disconnect” salute        his acknowledgement will indicate his acceptance of aircraft,        spacecraft and runway surface as suitable for the launch mission        subject to tower takeoff clearance, and he will change frequency        accordingly. The ground crew will remain clear of the taxiway        but in the general area until the takeoff is complete.    -   5. Special procedures will govern LB-1 flight operations,        including more restrictive takeoff weather minimums for ceiling,        visibility, crosswind, runway clutter and precipitation. Also        rejected takeoff, fuel dumping, primary and alternate launch        point (LP) criteria, will be more critical, especially        tropopause weather, particularly winds, which can be in excess        of 200 knots and turbulence which may be extreme. Alternate        launch points (LPs)/altitudes will be filed for every mission.    -   6. Staging will generally be conventional for the launch type        being conducted, however all specifications, exceptions,        alternate launch points (LPs) and other advisories will be        included on the flight plan and updated automatically or upon        request.    -   7. In the event of a failure in a primary launch system or        component, a joint decision will be reached between the captain        and the launch coordinator as to whether a safe/successful        launch can be achieved with a standby system or component or        hand-flown maneuver, or whether the load should be returned to        base or jettisoned, and if either of the latter, whether carrier        aircraft fuel dumping or another safer course of action is        indicated.    -   8. Although air-launch has demonstrated an excellent safety        record in both manned and unmanned missions, payload insurance        continues a major driver of launch cost, therefore every effort        should be extended to design equipment and procedures to the        highest standards of safe operation.

What is claimed:
 1. A system for launching aerospace payloads,comprising: an unmanned lifting body spacecraft including a payloadcompartment in a forward section of said spacecraft, rockets clusteredin an aft section of said spacecraft, reaction control system (RCS)modules configured to control a flight path of said spacecraft, andassociated avionics hardware and software; chines comprising wye-shapedstabilizers, wherein each arm of the wye is canted outboard from thevertical, and wherein a horizontal arm of the wye is fitted withelevons; and a carrier aircraft configured to air-launch saidspacecraft.